Thermally initiated variable venting system for rocket motor

ABSTRACT

A thermally initiated variable venting system may comprise a first linear shape charge (LSC) coupled to a first sensor and a second LSC coupled to a second sensor. An upper apex of the second LSC may be disposed within a lower apex of the first LSC. The output of the system may vary depending on whether the event is fast cook-off (FCO) or slow cook-off (SCO).

FIELD

The present disclosure relates generally to rocket motors, and moreparticularly, to a system which is active before the firing of a rocketmotor and which senses heat around the rocket motor which might ignitethe rocket motor or cause it to explode.

BACKGROUND

When a rocket motor is subjected to temperatures which would be reachedin a fortuitous fuel fire, the solid fuel of the rocket motor may igniteor explode. Unless steps are taken, ignition will cause thrust and therocket motor will be propelled. Should this occur in an enclosed spacesuch as a hangar or on an airport or a flight deck, the resultant rocketmotor flight is quite dangerous and destructive. Likewise, an explosiondue to rocket motor heating may lead to significant destruction andhazards surrounding the device.

SUMMARY

A thermally initiated variable venting system is disclosed, comprising afirst linear shape charge (LSC) coupled to a first sensor, and a secondLSC coupled to a second sensor, wherein the first LSC overlaps thesecond LSC, the first sensor is configured to activate to ignite thefirst LSC in response to at least a portion of the first sensor reachinga first temperature, the second sensor is configured to activate toignite the second LSC in response to at least a portion of the secondsensor reaching a second temperature, the thermally initiated variableventing system configured to project a molten jet in response to atleast one of the first LSC being ignited by the first sensor and thesecond LSC being ignited by the second sensor.

In various embodiments, the molten jet is greater in response to thefirst LSC being ignited by the first sensor.

In various embodiments, the molten jet is configured to cut a slotthrough a motor case in response to the first LSC being ignited by thefirst sensor.

In various embodiments, the molten jet is configured to cut a trenchinto a motor case in response to the second LSC being ignited by thesecond sensor.

In various embodiments, at least one of a velocity, a pressure, and atemperature of the molten jet is greater in response to the first LSCbeing ignited by the first sensor than in response to the second LSCbeing ignited by the second sensor.

In various embodiments, the molten jet is generated by a detonation wavethat travels in a direction perpendicular to the molten jet.

In various embodiments, the first temperature is less than the secondtemperature.

In various embodiments, the first sensor comprises a slow cook off (SCO)sensor and the second sensor comprises a fast cook off (FCO) sensor.

In various embodiments, the first LSC comprises a sheath defining a maincharge cavity and an explosive charge material contained in the maincharge cavity.

In various embodiments, the sheath comprises a hollowed chevron-shapedcross-section that defines the main charge cavity, the chevron-shapedcross-section defining an upper apex and a lower apex, and an upper apexof the second LSC is disposed in the lower apex of the first LSC.

A rocket motor is disclosed, comprising a rocket motor case, a covercoupled to the rocket motor case and extending longitudinally along anouter surface of the rocket motor case, a channel disposed in the cover,an opening of the channel facing the rocket motor case, a first linearshape charge (LSC) coupled to a first sensor, at least a portion of thefirst LSC disposed in the channel, and a second LSC coupled to a secondsensor, at least a portion of the second LSC disposed in the channel,wherein the second LSC is disposed between the first LSC and the rocketmotor case, the first sensor is configured to activate to ignite thefirst LSC in response to at least a portion of the first sensor reachinga first temperature, the second sensor is configured to activate toignite the second LSC in response to at least a portion of the secondsensor reaching a second temperature, wherein a molten jet is projectedfrom the opening in the channel in response to at least one of the firstLSC being ignited by the first sensor and the second LSC being ignitedby the second sensor.

In various embodiments, the molten jet is configured to cut a slotthrough a motor case in response to the first LSC being ignited by thefirst sensor.

In various embodiments, the molten jet is configured to cut a trenchinto a motor case in response to the second LSC being ignited by thesecond sensor.

In various embodiments, at least one of a velocity, a pressure, and atemperature of the molten jet is greater in response to the first LSCbeing ignited by the first sensor than in response to the second LSCbeing ignited by the second sensor.

In various embodiments, the first temperature is less than the secondtemperature.

In various embodiments, the first sensor comprises a slow cook off (SCO)sensor and the second sensor comprises a fast cook off (FCO) sensor.

In various embodiments, the first LSC comprises a sheath defining a maincharge cavity and an explosive charge material contained in the maincharge cavity.

In various embodiments, the sheath comprises a hollowed chevron-shapedcross-section that defines the main charge cavity, the chevron-shapedcross-section defining an upper apex and a lower apex, and an upper apexof the second LSC is disposed in the lower apex of the first LSC.

A method for manufacturing a thermally initiated variable venting systemis disclosed, comprising disposing a first linear shape charge (LSC)into a channel of a cover, disposing a second LSC into the channel, andcoupling the cover to a rocket motor case, wherein an opening in thechannel faces the rocket motor case, the second LSC disposed between thefirst LSC and the rocket motor case.

In various embodiments, the method further comprises disposing an upperapex of the second LSC within a lower apex of the first LSC.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand drawings are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter of the present disclosure is particularly pointed outand distinctly claimed in the concluding portion of the specification. Amore complete understanding of the present disclosure, however, may bestbe obtained by referring to the detailed description and claims whenconsidered in connection with the drawing figures.

FIG. 1A illustrates a schematic view of a solid rocket motor comprisinga thermally initiated variable venting system, in accordance withvarious embodiments;

FIG. 1B illustrates a cross section view of the thermally initiatedvariable venting system of FIG. 1A, in accordance with variousembodiments;

FIG. 2 illustrates a cross section view of a linear shaped charge, inaccordance with various embodiments;

FIG. 3 illustrates a cross section view of the thermally initiatedvariable venting system of FIG. 1A, schematically illustrating a moltenjet projected from a bottom linear shaped charge in response to thebottom linear shaped charge being ignited before a top linear shapedcharge, in accordance with various embodiments;

FIG. 4 illustrates a trench formed into a rocket motor case in responseto the molten jet of FIG. 3, in accordance with various embodiments;

FIG. 5 illustrates a cross section view of the thermally initiatedvariable venting system of FIG. 1A, schematically illustrating a moltenjet projected from a top linear shaped charge and a bottom linear shapedcharge in response to the top linear shaped charge being ignited beforethe bottom linear shaped charge, in accordance with various embodiments;

FIG. 6 illustrates a slot formed into a rocket motor case in response tothe molten jet of FIG. 5, in accordance with various embodiments;

FIG. 7 and FIG. 8 illustrate a section view of a solid rocket motorcomprising a thermally initiated variable venting system, in accordancewith various embodiments; and

FIG. 9 illustrates a method for manufacturing a thermally initiatedvariable venting system, in accordance with various embodiments.

DETAILED DESCRIPTION

The detailed description of various embodiments herein makes referenceto the accompanying drawings, which show various embodiments by way ofillustration. While these various embodiments are described insufficient detail to enable those skilled in the art to practice thedisclosure, it should be understood that other embodiments may berealized and that logical, chemical, and mechanical changes may be madewithout departing from the scope of the disclosure. Thus, the detaileddescription herein is presented for purposes of illustration only andnot of limitation. For example, the steps recited in any of the methodor process descriptions may be executed in any order and are notnecessarily limited to the order presented.

Furthermore, any reference to singular includes plural embodiments, andany reference to more than one component or step may include a singularembodiment or step. Also, any reference to attached, fixed, connected,or the like may include permanent, removable, temporary, partial, full,and/or any other possible attachment option. Additionally, any referenceto without contact (or similar phrases) may also include reduced contactor minimal contact.

With reference to FIG. 1A, a thermally initiated variable venting system(system) 100 is schematically illustrated, in accordance with variousembodiments. System 100 may include a rocket motor 110. Rocket motor 110may comprise a rocket motor case 112. Rocket motor case 112 may be inthe form of a cylindrical tube throughout most of its length. In variousembodiments, rocket motor case 112 may be the outer wall of the aft partof an airframe or may be a “slip-in” motor case housed within theairframe. Furthermore, the concepts of the present disclosure may beuseful for any system which utilizes a rocket motor. Rocket motor case112 may be normally closed, except for a nozzle 114 at its rear end.Rocket motor case 112 may carry a grain of solid rocket fuel therein.Upon combustion, this grain produces hot gas which raises the pressurewithin the rocket motor case. The hot gas is expelled from the nozzle114, producing thrust. Rocket motor case 112 is sufficiently strong towithstand the pressure of hot gas generation therein. The safety systemof the present disclosure is directed to opening the side of the rocketmotor case 112 or weakening it sufficiently so that internal hot gaspressure causes opening of the side as the pressure rises.

In various embodiments, a cover 120 may be coupled to rocket motor case112. Cover 120 may extend longitudinally along an outer surface of therocket motor case 112. In various embodiments, cover 120 may behermetically sealed to rocket motor case 112. A slow cook off (SCO)sensor 130 may be disposed in cover 120. A first linear shaped charge(LSC) 150 may extend from SCO sensor 130. Furthermore, a fast cook off(FCO) sensor 140 may be disposed in cover 120. A second LSC 160 mayextend from FCO sensor 140. The first LSC 150 may be disposed radiallyoutward from the second LSC 160. The first LSC 150 may overlap thesecond LSC 160. Stated differently, the first LSC 150 may be disposedover the second LSC 160. Stated yet differently, the second LSC 160 maybe disposed between the first LSC 150 and the rocket motor case 112.

In various embodiments, the SCO sensor 130 may be disposed oppositefirst LSC 150 from the FCO sensor 140. Similarly, the FCO sensor 140 maybe disposed opposite second LSC 160 from the SCO sensor 130. First LSC150 may be ignited by SCO sensor 130 at the end of first LSC 150 whichis coupled to SCO sensor 130. Second LSC 160 may be ignited by FCOsensor 140 at the end of second LSC 160 which is coupled to FCO sensor140. In this regard, first LSC 150 and second LSC 160 are ignited fromopposite ends. Furthermore, SCO sensor 130 and FCO sensor 140 mayoperate independently from each other.

In various embodiments, SCO sensor 130 may be configured to activate toignite first LSC 150 in response to at least a portion of SCO sensor 130reaching a first temperature (also referred to herein as an SCOtemperature) at which the propellant grain inside of the rocket motorcase 112 may be susceptible to being undesirably ignited. In variousembodiments, FCO sensor 140 may be configured to activate to ignitesecond LSC 160 in response to at least a portion of FCO sensor 140reaching a second temperature (also referred to herein as an FCOtemperature) at which the propellant grain inside of the rocket motorcase 112 may be susceptible to being undesirably ignited. Both SCOsensor 130 and FCO sensor 140 may respond to elevated environmentaltemperatures. Due to the difference in the manner in which the sensors(i.e., SCO sensor 130 and FCO sensor 140) detect temperature changesand/or rates of temperate change, the sensors may react in response todifferent types of stimuli.

In various embodiments, the SCO temperature is less than the FCOtemperature. In various embodiments, the FCO temperature is an externaltemperature (i.e., a temperature of the surrounding atmosphere) such asambient temperature. In various embodiments, the SCO temperature is aninternal temperature (i.e., a temperature of at least a portion of SCOsensor 130. In this manner, it is possible for either SCO sensor 130 orFCO sensor 140 to be activated before the other, even though the SCOtemperature may be less than the FCO temperature. For example, the FCOsensor 140 may detect an ambient temperature which is equal to orgreater than the FCO temperature before an internal temperature of theSCO sensor 130 reaches the SCO temperature, and vice versa.

With reference to FIG. 1B, cover 120 may comprise a trench or channel122 having an opening that faces rocket motor case 112. First LSC 150and second LSC 160 may be disposed in channel 122. First LSC 150 mayinclude a sheath 152 defining a main charge cavity 158. The sheath 152extends along a longitudinal axis (e.g., an X-axis) between a first endand a second end to define a sheath length (L_(S)) of the sheath 152.The sheath 152 may have a first hollowed chevron-shaped cross-sectionthat defines the main charge cavity 158. Second LSC 160 may similarlyinclude a sheath 162 defining a main charge cavity 168. With additionalreference to FIG. 2, which illustrates an isolated section view of firstLSC 150, the chevron-shaped cross-section may define an upper apex 181,and a lower apex 182. The upper apex of second LSC 160 may be located inthe lower apex of first LSC 150. The sheath 152 may be formed fromvarious materials including, but not limited to, aluminum, copper,tungsten, tantalum, lead, tin, cadmium, cobalt, magnesium, titanium,zinc, zirconium, molybdenum, beryllium, nickel, silver, gold, andplatinum. In various embodiments, second LSC 160 is similar to first LSC150.

With reference to FIG. 1B, first LSC 150 may further include anexplosive charge material (also referred to herein as a first explosivecharge material) 154 contained in the main charge cavity 158. Second LSC160 may further include an explosive charge material (also referred toherein as a second explosive charge material) 164 contained in the maincharge cavity 168. With main charge cavity 158 and main charge cavity168 filled with a respective explosive charge material 154, 164, firstLSC 150 and/or second LSC 160 are configured to generate a detonationwave which travels along the length (e.g., the X-direction) of first LSC150 and second LSC 160 which in turn projects a molten jet that travelsin a direction perpendicular to the detonation wave (e.g., the negativeZ-direction). The molten jet may be projected from the lower apex of theLSC.

With reference to FIG. 1A through FIG. 4, and with particular focus onFIG. 3 and FIG. 4, in response to second LSC 160 being ignited, thesecond LSC 160 may generate a detonation wave, which in turn projects amolten jet 302 that travels in a direction perpendicular to thedetonation wave (i.e., the negative Z-direction) and toward rocket motorcase 112. In this scenario, where the second LSC 160 is ignited first,the first LSC 150 may have minimal impact on the molten jet 302 and themolten jet 302 may merely score the rocket motor case 112. Stateddifferently, the molten jet 302 may partially cut through thewall-thickness of rocket motor case 112, creating a trench 480 (see FIG.4) therein.

With reference to FIG. 1A through FIG. 6, and with particular focus onFIG. 5 and FIG. 6, in response to first LSC 150 being ignited, the firstLSC 150 may generate a detonation wave, which in turn projects a moltenjet 502 that travels in a direction perpendicular to the detonation wave(i.e., the negative Z-direction) and towards second LSC 160, which inturn ignites second LSC 160 and generates an even greater molten jet504, which in turn is directed towards rocket motor case 112. Stateddifferently, the velocity, the pressure, and/or the temperature ofmolten jet 504 may be greater than that of molten jet 302 (see FIG. 3).In this scenario, where the first LSC 150 is ignited first, the effectof the first LSC 150 and the second LSC 160 are additively combined togenerate the greater (i.e., having a greater velocity, pressure, and/ortemperature than that of molten jet 302 of FIG. 3) molten jet 504 thanif the second LSC 160 had been ignited first. In this manner, molten jet504 may open the rocket motor case 112. Stated differently, the moltenjet 504 may cut completely through the wall-thickness of rocket motorcase 112, creating a slot 680 (see FIG. 6) therein. In this manner, thecutting performance of system 100 may vary depending on which LSC (i.e.,first LSC 150 and second LSC 160) is ignited first.

Referring to FIG. 1A through FIG. 1B, the main charge cavity 158 may befilled with a first type of explosive charge material 154 and the maincharge cavity 168 may be filled with a second type of explosive chargematerial 164. In various embodiments, explosive charge material 154 andexplosive charge material 164 may be the same type of explosive chargematerial, such as a Hexanitrostilbene (HNS) or a plastic-bondedexplosive (PBX), for example. In various embodiments, explosive chargematerial 154 may be different from explosive charge material 164. Invarious embodiments, explosive charge material 154 may be aplastic-bonded explosive (PBX) and explosive charge material 164 may bea Hexanitrostilbene (HNS). Upon detonation, the first explosive chargematerial and/or the second explosive charge material may produce adetonation wave having a detonation velocity. The detonation velocity ofthe explosive charge material dictates the rate at which the respectivedetonation wave propagates (i.e., the propagation rate).

In various embodiments, explosive charge material 154 may be packed at adifferent density than explosive charge material 164. In variousembodiments, the propagation rate of explosive charge material 154and/or explosive charge material 164 may be varied in response to thedensity thereof. In this regard, explosive charge material 154 may bepacked at a greater density than explosive charge material 164 toincrease the propagation rate of explosive charge material 154 relativeto explosive charge material 164. In at least one embodiment, thepacking density of the explosive charge material 154 may be greater thanthe packing density of the explosive charge material 164 by a ratioranging from approximately 1.1:1.0 to approximately 2.0:1.0. It isappreciated, however, that the packing density ratio is not limitedthereto. Furthermore, the propagation rate of explosive charge material154 and/or explosive charge material 164 may be varied by usingdifferent explosive charge materials having inherently differentpropagation rates.

In various embodiments, the performance of system 100 may be varied byincreasing and/or decreasing a cross-section area of main charge cavity158 and/or a cross-section area of main charge cavity 168. For example,the cross-section area of main charge cavity 158 and/or main chargecavity 168 may be varied to increase and/or decrease the cuttingperformance of molten jet 504 (see FIG. 5) depending, for example, onthe design (e.g., material and wall-thickness) of rocket motor case 112.Further, the cross-section area of main charge cavity 168 may be variedto increase and/or decrease the scoring performance of molten jet 302(see FIG. 3) depending, for example, on the design (e.g., material andwall-thickness) of rocket motor case 112.

Turning now to FIG. 7 and FIG. 8, a perspective section view of athermally initiated variable venting system (system) 700 is illustrated,in accordance with various embodiments. System 700 may be similar tosystem 100 (see FIG. 1A through FIG. 6). System 100 may include a rocketmotor 710 comprising a rocket motor case 712, and a cover 720 housing afirst LSC 750 extending from an SCO sensor 730 and a second LSC 760extending from an FCO sensor 740.

In various embodiments, at least a portion of FCO sensor 740 may beexposed to the atmosphere. The FCO sensor 740 may be configured toignite second LSC 760 in response to a preselected temperature andtemperature duration having been reached. For example, one or morethermal cords 802 may self-ignite in a maximum time of 30 seconds whenexposed to temperatures above 550° F. (288° C.). In various embodiments,one or more thermal cords 802 may self-ignite in a maximum time of 30seconds when exposed to temperatures above 600° F. (316° C.). Thermalcord 802 may be a pyrotechnic device which is specifically sensitive totemperature and is formulated to ignite (and provide a signal indicativethereof) in response to a preselected temperature and temperatureduration having been reached. The signal may be provided by ignition ofthe thermal cord 802 in a time less than the fast cook off time of therocket motor grain or other device being protected. The fast cook-offtime is the time that the motor is exposed to a given temperature with arequirement for survival (i.e. no explosion or ignition of the motorfuel grain). FCO sensor 740 may ignite LSC 760 using additionalenergetics and/or components disposed within FCO sensor 740.

When the second LSC 760 is ignited, it preferably cuts one or morestress raising notches in the outer portion of the rocket motor case712, or may cut directly through the rocket motor case 712 to the grain(see FIG. 4 and FIG. 6). The stress raising notches may be cut inselected locations along the length of the rocket motor case 712. Thesenotches or cuts are sufficient so that when the grain ignites, therocket motor case 712 splits and pressure is vented out of the splitside rather than developing pressure which causes significant thrust byexhausting from the nozzle. In this way, the rocket motor 710 isprevented from uncontrolled flight due to fire while the rocket motor isin storage, transport, or on an airplane prior to flight.

In various embodiments, SCO sensor 730 may comprise an intermetallicthermal sensor. SCO sensor 730 may ignite LSC 750 using additionalenergetics and/or components disposed within SCO sensor 730. In variousembodiments, the trigger temperature (SCO temperature) may range fromapproximately 290° F. to approximately 400° F. (143° C.-204° C.).

With reference to FIG. 9, a flow chart showing a method 900 for methodfor manufacturing a thermally initiated variable venting system isillustrated, in accordance with various embodiments. Method 900 includesdisposing a first linear shape charge (LSC) into a channel of a cover(step 910). Method 900 includes disposing a second LSC into the channel(step 920). Method 900 includes coupling the cover to the rocket motor(step 930).

With combined reference to FIG. 1B and FIG. 9, step 910 may includedisposing first LSC 150 into channel 122 of cover 120. Step 920 mayinclude disposing second LSC 160 into channel 122. In variousembodiments, step 920 may include disposing an upper apex (e.g., upperapex 181 of FIG. 2) of second LSC 160 within a lower apex (e.g., lowerapex 182 of FIG. 2)(see FIG. 2) of first LSC 150. Step 930 may includecoupling cover 120 to rocket motor case 112. Cover 120 may be coupled torocket motor case 112 such that opening 123 of channel 122 faces rocketmotor case 112. In various embodiments, cover 120 is coupled to rocketmotor case 112 via fasteners. In various embodiments, cover 120 iscoupled to rocket motor case 112 via an adhesive. In variousembodiments, cover 120 is coupled to rocket motor case 112 via a metaljoining process such as welding, soldering, brazing, or the like. It isappreciated, however, that cover 120 may be coupled to rocket motor case112 via any suitable method.

Benefits, other advantages, and solutions to problems have beendescribed herein with regard to specific embodiments. Furthermore, theconnecting lines shown in the various figures contained herein areintended to represent exemplary functional relationships and/or physicalcouplings between the various elements. It should be noted that manyalternative or additional functional relationships or physicalconnections may be present in a practical system. However, the benefits,advantages, solutions to problems, and any elements that may cause anybenefit, advantage, or solution to occur or become more pronounced arenot to be construed as critical, required, or essential features orelements of the disclosure. The scope of the disclosure is accordinglyto be limited by nothing other than the appended claims, in whichreference to an element in the singular is not intended to mean “one andonly one” unless explicitly so stated, but rather “one or more.”Moreover, where a phrase similar to “at least one of A, B, or C” is usedin the claims, it is intended that the phrase be interpreted to meanthat A alone may be present in an embodiment, B alone may be present inan embodiment, C alone may be present in an embodiment, or that anycombination of the elements A, B and C may be present in a singleembodiment; for example, A and B, A and C, B and C, or A and B and C.Different cross-hatching is used throughout the figures to denotedifferent parts but not necessarily to denote the same or differentmaterials.

Systems, methods and apparatus are provided herein. In the detaileddescription herein, references to “one embodiment”, “an embodiment”,“various embodiments”, etc., indicate that the embodiment described mayinclude a particular feature, structure, or characteristic, but everyembodiment may not necessarily include the particular feature,structure, or characteristic. Moreover, such phrases are not necessarilyreferring to the same embodiment. Further, when a particular feature,structure, or characteristic is described in connection with anembodiment, it is submitted that it is within the knowledge of oneskilled in the art to affect such feature, structure, or characteristicin connection with other embodiments whether or not explicitlydescribed. After reading the description, it will be apparent to oneskilled in the relevant art(s) how to implement the disclosure inalternative embodiments.

Furthermore, no element, component, or method step in the presentdisclosure is intended to be dedicated to the public regardless ofwhether the element, component, or method step is explicitly recited inthe claims. No claim element herein is to invoke 35 U.S.C. 112(f) unlessthe element is expressly recited using the phrase “means for.” As usedherein, the terms “comprises”, “comprising”, or any other variationthereof, are intended to cover a non-exclusive inclusion, such that aprocess, method, article, or apparatus that comprises a list of elementsdoes not include only those elements but may include other elements notexpressly listed or inherent to such process, method, article, orapparatus.

What is claimed is:
 1. A thermally initiated variable venting system fora rocket motor case, comprising: a first linear shape charge (LSC)coupled to a first sensor; and a second LSC coupled to a second sensor;wherein the first LSC radially overlaps the second LSC; the first sensoris configured to activate to ignite the first LSC in response to atleast a portion of the first sensor reaching a first temperature; thesecond sensor is configured to activate to ignite the second LSC inresponse to at least a portion of the second sensor reaching a secondtemperature; the thermally initiated variable venting system configuredto project a molten jet towards the rocket motor case in response to atleast one of the first LSC being ignited by the first sensor and thesecond LSC being ignited by the second sensor.
 2. The thermallyinitiated variable venting system of claim 1, wherein the molten jet isgreater in response to the first LSC being ignited by the first sensor.3. The thermally initiated variable venting system of claim 2, whereinthe molten jet is configured to cut a slot through a motor case inresponse to the first LSC being ignited by the first sensor.
 4. Thethermally initiated variable venting system of claim 2, wherein themolten jet is configured to cut a trench into a motor case in responseto the second LSC being ignited by the second sensor.
 5. The thermallyinitiated variable venting system of claim 2, wherein at least one of avelocity, a pressure, and a temperature of the molten jet is greater inresponse to the first LSC being ignited by the first sensor than inresponse to the second LSC being ignited by the second sensor.
 6. Thethermally initiated variable venting system of claim 1, wherein themolten jet is generated by a detonation wave that travels in a directionperpendicular to the molten jet.
 7. The thermally initiated variableventing system of claim 1, wherein the first temperature is less thanthe second temperature.
 8. The thermally initiated variable ventingsystem of claim 1, wherein the first sensor comprises a slow cook off(SCO) sensor and the second sensor comprises a fast cook off (FCO)sensor.
 9. The thermally initiated variable venting system of claim 1,wherein the first LSC comprises: a sheath defining a main charge cavity;and an explosive charge material contained in the main charge cavity.10. The thermally initiated variable venting system of claim 9, whereinthe sheath comprises a hollowed chevron-shaped cross-section thatdefines the main charge cavity, the chevron-shaped cross-sectiondefining an upper apex and a lower apex, and an upper apex of the secondLSC is disposed in the lower apex of the first LSC.
 11. A rocket motor,comprising: a rocket motor case; a cover coupled to the rocket motorcase and extending longitudinally along an outer surface of the rocketmotor case; a channel disposed in the cover, an opening of the channelfacing the rocket motor case; a first linear shape charge (LSC) coupledto a first sensor, at least a portion of the first LSC disposed in thechannel; and a second LSC coupled to a second sensor, at least a portionof the second LSC disposed in the channel; wherein the second LSC isdisposed between the first LSC and the rocket motor case; the firstsensor is configured to activate to ignite the first LSC in response toat least a portion of the first sensor reaching a first temperature; thesecond sensor is configured to activate to ignite the second LSC inresponse to at least a portion of the second sensor reaching a secondtemperature; wherein a molten jet is projected from the opening in thechannel in response to at least one of the first LSC being ignited bythe first sensor and the second LSC being ignited by the second sensor.12. The rocket motor of claim 11, wherein the molten jet is configuredto cut a slot through a motor case in response to the first LSC beingignited by the first sensor.
 13. The rocket motor of claim 11, whereinthe molten jet is configured to cut a trench into a motor case inresponse to the second LSC being ignited by the second sensor.
 14. Therocket motor of claim 11, wherein at least one of a velocity, apressure, and a temperature of the molten jet is greater in response tothe first LSC being ignited by the first sensor than in response to thesecond LSC being ignited by the second sensor.
 15. The rocket motor ofclaim 11, wherein the first temperature is less than the secondtemperature.
 16. The rocket motor of claim 11, wherein the first sensorcomprises a slow cook off (SCO) sensor and the second sensor comprises afast cook off (FCO) sensor.
 17. The rocket motor of claim 11, whereinthe first LSC comprises: a sheath defining a main charge cavity; and anexplosive charge material contained in the main charge cavity.
 18. Therocket motor of claim 17, wherein the sheath comprises a hollowedchevron-shaped cross-section that defines the main charge cavity, thechevron-shaped cross-section defining an upper apex and a lower apex,and an upper apex of the second LSC is disposed in the lower apex of thefirst LSC.
 19. A method for manufacturing a thermally initiated variableventing system, comprising: disposing a first linear shape charge (LSC)into a channel of a cover; disposing a second LSC into the channel;coupling a second sensor to the second LSC configured to activate toignite the second LSC in response to at least a portion of the secondsensor reaching a second temperature; and coupling the cover to a rocketmotor case, wherein an opening in the channel faces the rocket motorcase, the second LSC disposed between the first LSC and the rocket motorcase, such that a molten jet is projected from the opening towards therocket motor case in response to at least one of the first LSC beingignited by the first sensor and the second LSC being ignited by thesecond sensor.
 20. The method of claim 19, further comprising disposingan upper apex of the second LSC within a lower apex of the first LSC.